1: On SRM-1, a exit cone cover was installed on it prelaunch. What was the purpose of this cover?2: Did SRM-2 nozzle have any thermal cover on it like the SRM-1 nozzle?3: Was there any cover over the top of the ESS?

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"For Sardines, space is no problem!"-1996 Astronaut class slogan

"We're rolling in the wrong direction but for the right reasons"-USA engineer about the rollback of Discovery prior to the STS-114 Return To Flight mission

17 seconds, according to an old user manual. It took place immediately after staging. The Stage 2 burn did not start until about 2 minutes later.

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3: Specific impulse and thrust of the REM thrusters?

I haven't found a listing, but I did find that the post Stage 2 vernier burn lasted 6 m 40 sec and used 94.5 pounds of propellant (using multiple REM thrusters). The thrusters used hydrazine (N2H4) mono-propellant. Typical ISP for such thrusters is 220-230 seconds, give or take. I found one mention of Kaiser-Marquardt as the thruster manufacturer, but no details.

I haven't found a listing, but I did find that the post Stage 2 vernier burn lasted 6 m 40 sec and used 94.5 pounds of propellant (using multiple REM thrusters). The thrusters used hydrazine (N2H4) mono-propellant.

I found it interesting the motors were undersized to make sure that there was no over burn and the vernier burn made up for the intended shortfall and any additional shortfall. If there was a hot motor then the vernier burn was much shorter.

I haven't found a listing, but I did find that the post Stage 2 vernier burn lasted 6 m 40 sec and used 94.5 pounds of propellant (using multiple REM thrusters). The thrusters used hydrazine (N2H4) mono-propellant.

I found it interesting the motors were undersized to make sure that there was no over burn and the vernier burn made up for the intended shortfall and any additional shortfall. If there was a hot motor then the vernier burn was much shorter.

I suppose that injection accuracy was extra critical for direct to GEO missions (compared to, say, a non-adjusted Star 48 burn to a transfer orbit on a Delta II mission). Note also that the number of hydrazine tanks for the thrusters and RCS could be varied depending on the mission.

Based on a couple of early 1980s users guides, neither of which spell out this dimension specifically, my best estimate is that the 92 inch diameter cylindrical interstage part was nearly 70 inches long (I currently guess 69.89 inches) and was topped by a flared conical section that was 11.6 inches tall and 100 inches diameter at its top. Note that an extension of this conical flared section (20.2 inches tall and flaring out to 114 inches diameter) was part of Stage 2. It housed the Equipment Support Section.

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2: Extension time of the SRM-2 EEC?

17 seconds, according to an old user manual. It took place immediately after staging. The Stage 2 burn did not start until about 2 minutes later.

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3: Specific impulse and thrust of the REM thrusters?

I haven't found a listing, but I did find that the post Stage 2 vernier burn lasted 6 m 40 sec and used 94.5 pounds of propellant (using multiple REM thrusters). The thrusters used hydrazine (N2H4) mono-propellant. Typical ISP for such thrusters is 220-230 seconds, give or take. I found one mention of Kaiser-Marquardt as the thruster manufacturer, but no details.

Thanks Ed! Just to make sure I have gotten things right: There's a ring at the top of the conical section of the interstage and it's that one that is 100" in diameter? Or is it the where the conical section meets this ring?

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"For Sardines, space is no problem!"-1996 Astronaut class slogan

"We're rolling in the wrong direction but for the right reasons"-USA engineer about the rollback of Discovery prior to the STS-114 Return To Flight mission

Thanks Ed! Just to make sure I have gotten things right: There's a ring at the top of the conical section of the interstage and it's that one that is 100" in diameter? Or is it the where the conical section meets this ring?

Here's a drawing that shows how the conical (flared) section split into two parts, the top part was part of the second stage and the bottom part was part of the interstage (which itself remained attached to the first stage). The widest part, at the top, the "spacecraft interface plane", was 114 inches in diameter. The diameter at the stage separation plane, where the two IUS stages separated in the middle of the flared section, was 100 inches by my best estimate. The diameter at the bottom of the conical section was 92 inches, which was also the diameter of the cylindrical interstage section.

Here's a beautiful image of IUS-15/TDRS-E (5) during STS-43. A close look at the ring shows that the outer mold line of the trunnions and stiffeners, etc., doesn't run exactly parallel to the inner surface of the conical (flared) section. My 100 inch estimate was based on a straight line running between a 114 inch top diameter and a 92 inch bottom diameter at the top of the cylinder. Thus, the ring you point to could be a bit bigger than my estimate. We really need to see some detailed Boeing drawings, or get our hands on an IUS, to know for sure. There are one or two IUS mock ups out there in museums. See, for example, http://www.museumofflight.org/category/collections-locations/restoration-center-everett

Keep in mind that there were small, detail differences between Shuttle and Titan launched IUS, and possibly even between Titan 34D and Titan 4 versions. Note that the IUS exhibit in Seattle doesn't have this ring, which may have been for STS only. http://www.collectspace.com/images/news-110912c-lg.jpg

I'm also adding a blurry image of an IUS second stage, which shows the second, inner, hidden conical section.

Might anyone have on hand an old copy of an IUS/TDRS (or IUS/anything) "Cargo Systems Manual"? Those documents usually had detailed dimensional drawings of Shuttle payloads.

It includes some detail of the Stage 1/2 separation plane (STA 359.0). It seems to show a roughly 102 inch flat surface diameter and a roughly 111 inch maximum outer diameter at the ring. The extreme top part of the ESS is about 110 inches in diameter at the flat surface, and thus would presumably extend out to the 114 inch OD previously mentioned, if not further. A lot of detail surfaces on this thing!

Assuming that the IUS stages are used in such a way that the satellite is only carried to a standard GTO (as opposed to direct orbit insertion as it was done in history), what is the maximum payload that can be carried?